1. Field of the Invention
The present invention relates generally to a gas turbine engine, and more specifically to a turbine rotor disk assembly with a de-coupled platform.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
In a gas turbine engine, such as an aero or an industrial gas turbine engine, a compressor supplies a compressed air to a combustor, the combustor burns a fuel with the compressed air to produce a hot gas flow, and the hot gas flow is passed through a multiple staged turbine to extract mechanical power to drive the rotor shaft. In an aero engine, the rotor shaft is used to drive the compressor, while in an industrial gas turbine engine the rotor shaft drives the compressor and an external electric generator.
The industrial gas turbine engine is especially designed for the highest efficiency possible. Weight is not a major factor since the engine is secured in a stationary environment. The efficiency of a gas turbine engine can be increased by using a higher gas flow temperature passing into the turbine section. the gas flow temperature is limited to the material characteristics of the first stage turbine airfoils which include the stator vanes and the rotor blades.
To allow for a higher gas flow temperature in the first stage of the turbine section, improved cooling of the airfoils can be used. Turbine airfoils are designed with a complex arrangement of internal convection cooling passages and film cooling holes to maximize the airfoil cooling while minimizing the amount of pressurized cooling air used. Airfoil cooling circuits are customized in order to provide specific cooling amount over certain surfaces of the airfoils because not all of the surfaces are exposed to the same high gas flow temperatures.
In a rotor disk having a plurality of rotor blades extending from the disk and into the hot gas flow, each rotor blade is secured to the rotor disk through a slot, typically a fir tree shaped slot. In an industrial gas turbine engine, the size of the rotor blades is quite large. These large rotor blades also have a large mass. With a large mass in a rotating machine, the blades are exposed to high creep which can shorten the life of a rotor blade. In an industrial gas turbine engine, the engine runs for 24,000 to 48,000 hours before shutdown. Thus, the most efficient rotor blades are designed to have both light weight and resistance to high gas flow temperatures in order to provide for long life.
A prior art rotor blade includes an airfoil portion extending from a root portion with a platform formed at the lower airfoil portion to form an inner hot gas flow surface through the airfoil. The integral blade platform adds weight to the rotor blade. This extra weight on the rotor blade is carried by the blade root and the blade attachment slot in the rotor disk. Thus, the blade root must be designed to hold both the airfoil portion and the platform portion to the rotor disk.
A rotor blade made from a single crystal superalloy has a higher resistance to temperature than a non single crystal superalloy blade. However, forming a rotor blade with an integral platform from a single crystal material has major problems. One problem is that the platform extends substantially at a 90 degree angle from the single crystal direction of the rotor blade, which causes problems during casting. Many defective single crystal rotor blades are formed when the platform is formed integral to the blade.
Thus, it would be beneficial in the gas turbine engine to allow for a single crystal turbine blade to be used with a de-coupled platform for the purpose of allowing for the rotor blade to be made from a single crystal material to allow for higher gas flow temperatures. Also, it would be beneficial to form a rotor blade with a separate and de-coupled platform for the purpose of removing the loading due to the platform from the rotor blade root and slot attachment structure. In other words, the blade does not have to support the platform.
The U.S. Pat. No. 6,726,452 B2 issued to Strassberger et al on Apr. 27, 2004 and entitled TURBINE BLADE ARRANGEMENT discloses a turbine rotor blade assembly with adjacent rotor blades secured within the rotor disk slots and a platform un-coupled from the rotor blades and secured to the rotor disk by a holding device. The Strassberger invention un-couples the platform from the rotor blades and allows for a single crystal rotor blade, but has several problems in which the present invention solves. One problem with the Strassberger invention is that a large air gap is formed between the platform and the rotor disk that will allow for hot gas flow injection and require purge air to cool the space.
Also a problem in the prior art turbine rotor blades, the hot gas migration phenomenon on the airfoil pressure side is created by the combination of hot flow core gas axial velocity and static pressure gradient exerting on the surfaces of the airfoil pressure wall and the suction wall of an adjacent airfoil. Because of this hot gas flow, some of the hot core gas flow from the upper airfoil span is transferred toward a close proximity to the platform and thus creates a high heat transfer coefficient and high gas temperature region at approximately two-thirds of the blade chord location. FIG. 1 shows a cut-away view of the vortices formation of the hot flow gas migration across the turbine flow passage.
Cooling of the blade fillet region and platform by means of conventional backside convective cooling yields inefficient results due to the thickness of the airfoil fillet region and not being able to utilize effective cooling technique for the blade platform. As a result, a thermal mismatch between the blade airfoil and the platform creates LCF deficiency for the blade, and especially for a blade with a high mass platform.